This disclosure relates generally to gas turbine engines and, more particularly, to cooling techniques for the airfoil sections of turbine blades of the engine. In particular, the present application is directed to cooling techniques for blade platforms.
In general, gas turbine engines are built around a power core comprising a compressor, a combustor and a turbine, which are arranged in flow series with a forward (upstream) inlet and an aft (downstream) exhaust. The compressor compresses air from the inlet, which is mixed with fuel in the combustor and ignited to produce hot combustion gases. The hot combustion gases drive the turbine section, and are exhausted with the downstream flow.
The turbine drives the compressor via a shaft or a series of coaxially nested shaft spools, each driven at different pressures and speeds. The spools employ a number of stages comprised of alternating rotor blades and stator vanes. The vanes and blades typically have airfoil cross sections, in order to facilitate compression of the incoming air and extraction of rotational energy in the turbine. The blades are secured to the rotor disk through a blade platform.
High combustion temperatures also increase thermal and mechanical loads, particularly on turbine airfoils and associated platforms downstream of the combustor. This reduces service life and reliability, and increases operational costs associated with maintenance and repairs.
Blade platforms have been passively cooled by leakage air in a large plenum or a few filmholes, resulting in low backside heat transfer coefficients and high metal temps. Small cooling chambers are required to adequately cool the platform. However, these small chambers result in the feed holes that supply cooling air to these chambers being located in an area of the blade neck that is difficult to drill and has high stress due to the platform centrifugal loads.
Accordingly, it is desirable to provide cooling to the blade platforms in an efficient manner.